Efficient gas turbine engine installation and operation

ABSTRACT

A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

CROSS-REFERENCE TO RELATED APPLICATION(S)

This application is a continuation of U.S. application Ser. No.16/713,666 filed Dec. 13, 2019, which is a continuation of U.S.application Ser. No. 16/437,356 filed Jun. 11, 2019, which is based onand claims priority under 35 U.S.C. 119 from British Patent ApplicationNo. 1903262.2 filed on Mar. 11, 2019. The entire contents of the aboveapplications are incorporated herein by reference.

The present disclosure relates to a gas turbine engine. Aspects of thepresent disclosure relate to a gas turbine engine having a configurationoptimized for improved efficiency, installation and/or operability.

Modern gas turbine engines comprise a fan that is driven by a turbine.At least a part of the fan flow bypasses the core of the engine, andinstead flows through a bypass duct to produce thrust. The flow thatdoes pass into the core of the gas turbine engine is compressed in acompressor before being combusted, and then expanded through a turbine.

Optimizing the design of such gas turbine engines requires a number ofdifferent, often competing, factors to be balanced. For example, it isdesirable to be able to optimize the overall efficiency of the gasturbine engine installation with an aircraft, but there are a number ofdifferent elements that combine to produce the overall efficiency. Itwould be desirable to provide a gas turbine engine that providesimproved overall efficiency when installed with an airframe.Furthermore, it has been recognised that the pursuit of improvedefficiency must not come at the expense of other factors, such as engineoperability and maintenance requirements.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising a turbine, a compressor, and a combustor;a fan (which may be located upstream of the engine core) comprising aplurality of fan blades; anda gearbox that receives an input from a core shaft that is connected toat least a part of the turbine, the gearbox outputting drive to the fanso as to drive the fan at a lower rotational speed than the core shaft,wherein:a bypass duct is defined radially outside the core, with the leadingedge of a splitter defining the point at which flow splits into coreflow and bypass flow in use;a core compressor aspect ratio is defined as the ratio of the axialdistance between the leading edge of the splitter and the leading edgeof the tip of the most downstream compressor blade to the radius of theleading edge of the splitter;a core compressor pressure ratio, defined as the pressure immediatelydownstream of the final rotor blade in the compressor divided by thepressure immediately upstream of the first rotor blade in the corecompressor at cruise conditions, is in the range of from 34 to 60;the core compressor comprises twelve, thirteen or fourteen rotor stages;andthe ratio of the core compressor aspect ratio divided by the corecompressor pressure ratio is in the range of from 0.03 to 0.09.

It has been found that providing a gas turbine engine as claimed hereinresults in an engine that has a sufficiently high core compressorpressure ratio to achieve high thermal efficiency without requiring anexcessively long compressor for a given radius of the leading edge ofthe splitter (which may itself be set by other design considerations,including, for example, the outer diameter of the gearbox). This may beenabled at least in part by achieving the core compressor pressure ratiousing twelve, thirteen or fourteen rotor stages in the core compressor(i.e. excluding the fan). Fewer than twelve stages has been found toresult in poor compression efficiency, whereas greater than fourteenstages has been found to result in an excessively long compressor.Optionally, the core compressor pressure ratio is achieved using exactlytwelve (i.e. twelve and no more than twelve) rotor stages in the corecompressor.

Achieving a core compression ratio over a relatively short compressorprovides installation benefits when mounting the gas turbine engine toan aircraft—for example in terms of its interaction with other aircraftstructures, such as the wing, leading to reduced overall aircraft fuelburn and/or the ability to install the engine without compromising otheraircraft structures.

Still further, it has been found that geared gas turbine engines inparticular tend to suffer from “rotor bow”. This results fromdifferential cooling (in the vertical plane) of one or more core shaftsof the engine that connect the compressor to the turbine when the engineis shut down after use. This differential cooling—which is caused simplyby the tendency for the warm air to rise towards the upper side of theshaft—results in the core shaft bending (or “bowing”) as it cools. Ifremedial (and potentially costly and time-consuming) measures are nottaken, this can result in the core shaft becoming locked in position foran extended period of time until it is fully cooled, during which periodof time it is not possible to start the engine, and thus the aircraft towhich it is attached is grounded. It has been found that providing a gasturbine engine having a ratio of the core compressor aspect ratiodivided by the core compressor pressure ratio—as enabled by the claimednumber of core compressor rotor stages—in the ranges defined and/orclaimed herein provides a high thermal efficiency but with a greatlyreduced risk of rotor bow. If the ratio of the core compressor aspectratio divided by the core compressor pressure ratio is reduced stillfurther below the lower bound disclosed and/or claimed herein (and/orthe number of compression stages is reduced below twelve), it has beenfound that either the compression efficiency drops dramatically due tothe requirement of providing a relatively high compression ratio over adistance that is too short, or the engine become particularly sensitiveto rotor bow if the axial distance is increased in order to improvecompression efficiency.

Optionally, the ratio of the core compressor aspect ratio divided by thecore compressor pressure ratio is in the range having a lower bound ofany one of 0.04, 0.045 or 0.05 and an upper bound of any one of 0.06,0.07, 0.08 or 0.085.

The core compressor aspect ratio (CCAR) may be, for example, in therange having a lower bound of any of 1.7, 1.8, 1.9, 2, 2.1, 2.2 or 2.3and an upper bound of any one of 4.2, 4, 3.8, 3.6, 3.4, 3.2, 3, 2.9 or2.8, for example in the range of from 1.7 to 4.2; 1.8 to 3.4; 2.0 to2.9; 2.1 to 2.9; or 2.3 to 2.8.

Each rotor stage may be axially separated from its neighbouring rotorstages. A stator stage may be provided between each pair of neighbouringrotor stages.

According to a second aspect there is provided a gas turbine engine foran aircraft comprising:

an engine core comprising a turbine, a compressor, and a combustor;a fan (which may be located upstream of the engine core) comprising aplurality of fan blades; anda gearbox that receives an input from a core shaft that is connected toat least a part of the turbine, the gearbox outputting drive to the fanso as to drive the fan at a lower rotational speed than the core shaft,wherein:a compression system radius ratio defined as the ratio of the radius ofthe tip of a fan blade to the radius of the tip of the most downstreamcompressor blade is in the range of from 5 to 9.

The most downstream compressor blade (which would typically be any oneof a plurality of blades in a row) may be in the axially most downstreamcompressor blade row and/or a compressor blade in the blade row that isclosest to the combustor and/or immediately upstream of the combustor(that is, without any in-between blade rows) and/or in the highestpressure compressor blade row.

The radius of the tip of the fan blade may be defined as the radius atthe tip on the leading edge of the blade. Similarly, the radius of thetip of the most downstream compressor blade may be defined as the radiusat the tip on the leading edge of the blade.

In some arrangements, the compression system radius ratio (CSRR) may bein a range having a lower bound of any one of 5, 5.1, 5.2, 5.3, 5.4,5.5, 5.6, 5.7, 5.8, 5.9 or 6, and/or an upper bound of any one of 9,8.9, 8.8, 8.7, 8.6, 8.5, 8.4, 8.3, 8.2, 8.1, 8, 7.9, 7.8, 7.7, 7.6, 7.5,7.4, 7.3, 7.2, 7.1, 7, 6.9, 6.8, 6.7, 6.6 or 6.5. Purely by way ofexample, the CSRR may be in the range of from 5.2 to 8.5 or 5.3 to 7.2.

It has been found that gas turbine engines having arrangements in whichthe compression system radius ratio is in the ranges defined herein mayprovide improved overall efficiency of the gas turbine engine wheninstalled with an airframe. For example, the compression system radiusratio in the ranges defined herein may result in an optimized balancebetween high propulsive efficiency (which typically improves with,amongst other factors, a relatively high diameter fan), high thermalefficiency (which typically improves with, amongst other factors,increasing core pressure ratio) and improved installation qualities,including aerodynamic interaction with the airframe and improvedflexibility in positioning of the core (and thus the whole engine)relative to the airframe. In this regard, a low radius of the tip of themost downstream compressor blade relative to the radius of the fan blademay result in such improved flexibility in positioning the core, forexample allowing the centreline of the engine to be positionedvertically closer to the aircraft wing for a given fan diameter. In somecases this may increase the maximum fan diameter that it is possible toinstall on a given airframe, which may in turn provide a propulsiveefficiency advantage.

On the other hand, it has been found that reducing the radius of the tipof the most downstream compressor blade relative to the radius of thefan blade further, such that the CSRR is above the ranges defined herein(for example above 9), may be detrimental to the overall enginecharacteristics. This may be due to an intolerable decrease incompression efficiency (for example due to a higher than optimalcompression ratio over a given axial length) and/or the installationbenefits becoming compromised, for example due to increased compressorlength (and thus engine length) required in order to achieve acceptablecompression efficiency.

Still further, it has been found that providing a gas turbine enginehaving a CSRR in the ranges defined herein may provide high thermal andpropulsive efficiency but with a greatly reduced risk of rotor bow, asdescribed elsewhere herein. If the CSRR is increased still further abovethe upper bound disclosed herein, it has been found that either thecompression efficiency may drop due to the requirement of providing ahigh enough compression ratio to maintain acceptable thermal efficiencyover a distance that is too short, or the engine (and shaft) length mayincrease (in order to maintain acceptable compression efficiency) and itmay become more susceptible to rotor bow.

Accordingly, it has been found that a CSRR in the defined ranges mayprovide a gas turbine engine that has high operability/low maintenancerequirements, together with high efficiency when installed with anairframe, for example in terms of overall fuel burn and/or installationcapability.

The gearbox assists in enabling the gas turbine engine to achieve thedisclosed compression system radius ratio.

The compression system radius ratio divided by an engine core radiusratio may be in the range of from 5.5 to 10, for example 6 to 8, forexample in a range having a lower bound of any one of 5.5, 5.6, 5.7,5.8, 5.9, 6, 6.1, 6.2, 6.3, 6.4, 6.5 and/or an upper bound of any one of10, 9.5, 9, 8.5, 8, 7.9, 7.8, 7.7, 7.6, 7.5, 7.4, 7.3, 7.2, 7.1, or 7.In this regard, the engine core radius ratio (ECRR) is as definedelsewhere herein, i.e. the ratio of the radius of the tip of the mostdownstream turbine blade in the engine to the radius of the leading edgeof the splitter. Such arrangements may, in some cases, further improvethe overall efficiency and/or installation and/or maintenancerequirements of the gas turbine engine when installed with an airframe.

The compression system radius ratio divided by a core compressor aspectratio may be in the range of from 1.7 to 4.2, for example in a rangehaving a lower bound of any one of 1.7, 1.8, 1.9, 2, 2.1, 2.2, 2.3, 2.4,or 2.5 and/or an upper bound of any one of 4.2, 4.1, 4, 3.9, 3.8, 3.7,3.6, 3.5, 3.4, 3.3, 3.2, 3.1, 3, 2.9, 2.8 or 2.7, for example in therange of from 1.8 to 2.9. In this regard, the core compressor aspectratio (CCAR) is as defined elsewhere herein, i.e. as the ratio of theaxial distance between the leading edge of the splitter and the leadingedge of the tip of the most downstream compressor blade to the radius ofthe leading edge of the splitter. Such arrangements may, in some cases,further improve the overall efficiency and/or installation and/ormaintenance requirements of the gas turbine engine when installed withan airframe. For example, such arrangements may have a particularlycompact core compression system (for example in terms of axial length).

A compression system speed ratio may be as defined elsewhere herein,i.e. as the ratio of the rotational speed of the most downstreamcompressor blade to the rotational speed of the fan at cruiseconditions. The product of the compression system radius ratio and thecompression system speed ratio may be in the range of from 25 to 80, forexample in a range having a lower bound of any one of 25, 30, 35, 40, 45and/or an upper bound of any one of 80, 75, 70, 65, 60, 55 or 50. Sucharrangements may, in some cases, further improve the overall efficiencyand/or installation and/or maintenance requirements of the gas turbineengine when installed with an airframe.

Purely by way of example, of the radius of the tip of a fan blade may bein the range of from 120 cm to 140 cm and the radius of the tip of themost downstream compressor blade may be in the range of from 17 cm to 28cm. Purely by way of further example, of the radius of the tip of a fanblade may be in the range of from 165 cm to 190 cm and the radius of thetip of the most downstream compressor blade may be in the range of from24 cm to 35 cm.

According to a third aspect, there is provided a gas turbine engine foran aircraft comprising:

an engine core comprising a turbine, a compressor, and a combustor;a fan comprising a plurality of fan blades, each fan blade having a fanblade height defined as the radius of the leading edge at the tip of theblade minus the radius of the point where the leading edge intersectsthe radially inner gas-washed hub; anda gearbox that receives an input from a core shaft that is connected toat least a part of the turbine, the gearbox outputting drive to the fanso as to drive the fan at a lower rotational speed than the core shaft,wherein:the blade height of the most downstream compressor blade is defined asthe radius of the leading edge at the tip of the blade minus the radiusof the point where the leading edge intersects the radially innergas-washed surface; and a compression system blade ratio (CSBR) definedas the ratio of the fan blade height to the height of the mostdownstream compressor blade is in the range of from 45 to 95.

The compression system blade ratio may be in a range having a lowerbound of any one of 45, 46, 48, 50, 52, 54 or 56 and an upper bound ofany one of 95, 90, 85, 80, 75 or 70.

It has been found that gas turbine engines having arrangements in whichthe compression system blade ratio is in the ranges defined herein mayprovide improved overall efficiency of the gas turbine engine wheninstalled with an airframe. For example, the compression system bladeratio in the ranges defined herein may result in an optimized balancebetween high propulsive efficiency (which typically improves with,amongst other factors, a relatively large fan blade), high thermalefficiency (which typically improves with, amongst other factors,increasing core pressure ratio) and improved installation qualities,including aerodynamic interaction with the airframe and improvedflexibility in positioning of the core (and thus the whole engine)relative to the airframe.

In particular, high propulsive efficiency may be achieved by having ahigh bypass ratio which may be enabled by a relatively large fan bladeheight. High thermal efficiency may be achieved by having a highcompressor pressure ratio, which may be enabled by a relatively smallheight of the highest pressure compressor blade for a given fan bladeheight. Dropping the CSBR below the lower limit defined herein mayreduce the overall engine efficiency due to the combined effects of thethermal efficiency and propulsive efficiency. However, if the CSBR isfurther increased beyond the upper limits defined herein, this may havean adverse impact on other engine characteristics. For example, it hasbeen found that further increases in the CSBR may require thecompression system to become unacceptably long in order to achieveacceptable levels of compression efficiency. This may lead to adverseimpacts at an aircraft level, for example in terms of the engine'sinteraction with other aircraft structures, such as the wing, leading toreduced overall fuel burn and/or the ability to install the enginewithout compromising other aircraft structures.

Still further, it has been found that providing a gas turbine enginehaving a CSBR in the ranges defined herein may provide high thermal andpropulsive efficiency but with a greatly reduced risk of rotor bow(which is described elsewhere herein). If the CSBR is increased stillfurther above the upper bound defined herein, it has been found thateither the compression efficiency may drop due to the requirement ofproviding a high enough compression ratio to maintain acceptable thermalefficiency over a distance that is too short, or the engine (and shaft)length may increase (in order to maintain acceptable compressionefficiency), resulting in an increased susceptibility to rotor bow.

The gearbox assists in enabling the gas turbine engine to achieve thedisclosed CSBR.

The compression system blade ratio divided by an engine core radiusratio may be in the range of from 50 to 95, for example in a rangehaving a lower bound of 45, 50 or 52 and an upper bound of 65, 70, 75,80, 85, 90 or 95, optionally 50 to 85 or 50 to 75. In this regard, theengine core radius ratio (ECRR) is as defined elsewhere herein, i.e. theratio of the radius of the tip of the most downstream turbine blade inthe engine to the radius of the leading edge of the splitter. Sucharrangements may, in some cases, further improve the overall efficiencyand/or installation and/or maintenance requirements of the gas turbineengine when installed with an airframe.

The compression system blade ratio divided by a core compressor aspectratio may be in the range of from 15 to 50, for example in a rangehaving a lower bound of 15, 16, 17, 18, 19 or 20 and an upper bound of50, 45, 40, 35 or 30, optionally 16 to 40. In this regard, the corecompressor aspect ratio (CCAR) is as defined elsewhere herein, i.e. asthe ratio of the axial distance between the leading edge of the splitterand the leading edge of the tip of the most downstream compressor bladeto the radius of the leading edge of the splitter. Such arrangementsmay, in some cases, further improve the overall efficiency and/orinstallation and/or maintenance requirements of the gas turbine enginewhen installed with an airframe. For example, such arrangements may havea particularly compact core compression system (for example in terms ofaxial length).

A compression system speed ratio may be as defined elsewhere herein,i.e. as the ratio of the rotational speed of the most downstreamcompressor blade to the rotational speed of the fan at cruiseconditions. The product of the compression system blade ratio and thecompression system speed ratio may be in the range of from 300 to 800,optionally 320 to 750, optionally 325 to 700. Such arrangements may, insome cases, further improve the overall efficiency and/or installationand/or maintenance requirements of the gas turbine engine when installedwith an airframe.

According to a fourth aspect, there is provided a gas turbine engine foran aircraft comprising:

an engine core comprising a turbine, a compressor, and a combustor; anda fan comprising a plurality of fan blades, wherein:a compression system speed ratio, defined as the ratio of the rotationalspeed of the most downstream compressor blade to the rotational speed ofthe fan at cruise conditions, is in the range of from 6 to 10, forexample in a range having a lower bound of any one of 6, 6.1, 6.2, 6.3,6.4, 6.5, 6.6, 6.7, 6.8, 6.9, 7, 7.1, 7.2, 7.3, 7.4, 7.5 and/or an upperbound of any one of 10, 9.5, 9, 8.9, 8.8, 8.7, 8.6, 8.5, 8.4, 8.3, 8.2,8.1 or 8. By way of example, the compression system speed ratio may bein the range of from 7 to 9, for example 7.4 to 8.5, with the diameterof the fan optionally being in the range of from 320 to 390 cm and/orthe rotational speed of the fan at cruise conditions being in the rangeof from 1300 rpm to 1800 rpm. In such an arrangement, the diameter ofthe fan may be in the range of from 230 cm to 400 cm. In such anarrangement, the turbine may comprise a first turbine and a secondturbine; the compressor may comprise a first compressor and a secondcompressor. The first turbine and first compressor may be connected by afirst core shaft, and the second turbine and second compressor may beconnected by a second core shaft core. The second turbine, secondcompressor, and second core shaft may be arranged to rotate at a higherrotational speed than the first core shaft. In such an arrangement, themost downstream compressor blade would part of the second compressor.

Such a gas turbine engine according to the fourth aspect may comprise agearbox that receives an input from a core shaft that is connected to atleast a part of the turbine, the gearbox outputting drive to the fan soas to drive the fan at a lower rotational speed than the core shaft.

It will be appreciated that compatible features from any aspect may becombined. Purely by way of example, a gas turbine engine according to anaspect may comprise any one or more of the ranges disclosed herein for:CSRR; CCAR; ECRR; CSSR; CSBR; core compressor pressure ratio; CCAR/(corecompressor pressure ratio); CSRR/ECRR; CSRR/CCAR; CSRR*CSSR; CSBR/ECRR;CSBR/CCAR; CSBR*CSSR; number of compressor rotor stages.

Furthermore, the any one or more of the following features may apply toand/or be incorporated into any aspect of the present disclosure.

The first rotor blade in the core compressor may be referred to as thefirst rotor blade downstream of the fan. The final rotor blade in thecompressor may be referred to as the highest pressure compressor blade,or the first rotor blade upstream of the combustor. Optionally, the corecompressor pressure ratio (which is defined at cruise conditions) may bein the range having a lower bound of any one of 33, 34, 35, 36, 38 or 40and an upper bound of any one of 52, 55, 57 or 60, for example 36 to 52at cruise conditions. According to the definition provided herein, thecore compressor pressure ratio does not include the pressure rise due tothe fan.

A bypass duct is defined radially outside the core. The leading edge ofa splitter defines the point at which flow splits into core flow andbypass flow (core flow being flow that goes through the core of theengine in use, and bypass flow being flow that goes into the bypass ductin use). An engine core radius ratio (ECRR) may then be defined as theratio of the radius of the tip of the most downstream turbine blade inthe engine to the radius of the leading edge of the splitter. The ECRRmay be less than 1, for example in the range of from 0.75 to 1, 0.8 to0.98, 0.81 to 0.94, or 0.82 to 0.93. Such arrangements may, in somecases, further improve the overall efficiency and/or installation and/ormaintenance requirements of the gas turbine engine when installed withan airframe.

The fan blades may be surrounded by a nacelle, which may include a fancasing. Thus, the radially outer tips of the fan blades may besurrounded by a radially inner gas washed surface of the nacelle.Accordingly, the gas turbine engine may be referred to as a turbofan gasturbine engine and/or the fan may be referred to as a ducted fan. Thenacelle may form a radially outer surface of the bypass duct of such aturbofan engine.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.25 to 0.32, or 0.28 to 0.32.These ratios may commonly be referred to as the hub-to-tip ratio. Theradius at the hub and the radius at the tip may both be measured at theleading edge (or axially forwardmost) part of the blade. The hub-to-tipratio refers, of course, to the gas-washed portion of the fan blade,i.e. the portion radially outside any platform.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A gas turbine engine as described and/or claimed herein may furthercomprise an intake that extends upstream of the fan blades. An intakelength L may be defined as the axial distance between the leading edgeof the intake and the leading edge of the tip of the fan blades. The fandiameter D may be as defined elsewhere herein, i.e. the diameter of thefan at the leading edge of the tips of the fan blades. The ratio L/D maybe less than 0.5, for example in the range of from 0.2 to 0.45, 0.25 to0.4 or less than 0.4. Where the intake length varies around thecircumference, the intake length L used to determine the ratio of theintake length to the diameter D of the fan may be measured at the π/2 or3π/2 positions from top dead centre of the engine (i.e. at the 3 o'clockor 9 o'clock positions), or the average of the intake length at thesetwo positions where they are different.

Arrangements of the present disclosure may be particularly beneficialfor fans that are driven via a gearbox. The gearbox receives an inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft. The input to thegearbox may be directly from the core shaft, or indirectly from the coreshaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gearbox is a reduction gearbox (in that the output to the fan is alower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1,3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratiomay be, for example, between any two of the values in the previoussentence. Purely by way of example, the gearbox may be a “star” gearboxhaving a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theengine core may comprise a first turbine, connected to a firstcompressor by a first core shaft. The engine core may further comprise asecond turbine, a second compressor, and a second core shaft connectingthe second turbine to the second compressor. The second turbine, secondcompressor, and second core shaft may be arranged to rotate at a higherrotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor. In such anarrangement, the most downstream compressor blade would be a part of thesecond compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160inches) or 420 cm (around 165 inches). The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allunits in this paragraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypassduct may be substantially annular. The bypass duct may be radiallyoutside the core engine. The radially outer surface of the bypass ductmay be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressure atthe exit of the highest pressure compressor (before entry into thecombustor) to the stagnation pressure upstream of the fan. By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 45 to 70,or 50 to 65. Note that the overall pressure ratio differs from the corecompressor pressure ratio because the overall pressure ratio alsoincludes the pressure rise over the fan root (i.e. the portion of thefan over which air that subsequently flows into the engine core passes).

In some arrangements, a fan pressure ratio, defined as the ratio of themean total pressure of the flow at the fan exit to the mean totalpressure of the flow at the fan inlet, may be no greater than 1.5 atcruise conditions, for example no greater than 1.45, 1.4 or 1.35. Thefan pressure ratio may be in the range of from 1.35 to 1.43, for exampleon the order of 1.39.

Downstream of the fan, the flow through the gas turbine engine is splitinto a core flow (which flows through the engine core) and a bypass flow(which flows through the bypass duct). The gas turbine engine comprisesa splitter (which may be an annular splitter) at which the flow isdivided between the core flow that flows through the engine core, andthe bypass flow that flows along a bypass duct. In some arrangements, afan root pressure ratio, defined as the ratio of the mean total pressureof the flow at the fan exit that subsequently flows through the enginecore to the mean total pressure of the flow at the fan inlet, may be nogreater than 1.3 at cruise conditions.

In some arrangements, the fan root pressure ratio at cruise conditionsmay be no greater than 1.24, for example no greater than 1.23, forexample no greater than 1.22, for example no greater than 1.21, forexample no greater than 1.2. In some arrangements, the fan root pressureratio at cruise conditions may be in the range of from 1.18 to 1.30, forexample 1.21 to 1.27.

Where the term mean is used herein in relation to a pressure (forexample a total pressure), this may be (for example) an area averagetaken over the relevant surface.

A fan root to tip pressure ratio, defined as the ratio of the mean totalpressure of the flow at the fan exit that subsequently flows through theengine core to the mean total pressure of the flow at the fan exit thatsubsequently flows through the bypass duct, may be no greater than (forexample less than) 0.95, for example no greater than 0.94, 0.93, 0.92,0.91 or 0.9 at cruise conditions, for example in the range of from 0.8to 0.95, for example 0.82 to 0.89, for example 0.83 to 0.88.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹'s, 105 Nkg⁻¹'s, 100 Nkg⁻¹'s, 95 Nkg⁻¹'s, 90 Nkg⁻¹'s, 85 Nkg⁻¹'s or80 Nkg⁻¹'s. The specific thrust may be in an inclusive range bounded byany two of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹'s to100 Nkg⁻¹'s, or 85 Nkg⁻¹'s to 95 Nkg⁻¹'s. Such engines may beparticularly efficient in comparison with conventional gas turbineengines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein. Such an aircraft may comprise,for example, 2, 3 or 4 such gas turbine engines. At least one engine(for example 1, 2 or more than 2 engines) may be attached to each wingof such an aircraft. As mentioned elsewhere herein, the installedengines may contribute significantly to overall aircraft efficiency, forexample due to an improved combination of gas turbine engine efficiencytogether with advantages provided by improved installation with theairframe. Further operability and/or maintenance advantages may also beachieved, for example due to reduced susceptibility to rotor bow.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine in accordancewith an example of the present disclosure;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine in accordance with an example of the present disclosure;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine and

FIG. 4 is a schematic of a gas turbine engine in accordance with anexample of the present disclosure.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

Gas turbine engines are susceptible to a phenomenon known as “rotorbow”. As described elsewhere herein, this results from differentialcooling of one or more of the shafts 26, 27 when the engine is shut downafter use, and can result in the engine being inoperative for anextended period of time after shut down, at least in the absence oftime-consuming and/or expensive remedial action. It has been found thatthis problem may be exacerbated on modern engines, particularly thosewith a gearbox and/or high compression ratio. As explained elsewhereherein, the gas turbine engines 10 described and/or claimed herein mayhave a high efficiency (for example in terms of propulsive and/orthermal efficiency) but with a greatly reduced risk of rotor bowaffecting the shafts 26, 27.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

It will be appreciated that FIG. 1 is not necessarily to scale in allaspects, and is included merely to aid the description. FIG. 4 is aschematic representation of a gas turbine engine according to thepresent disclosure, and is provided to illustrate the dimensionsreferred to herein. Again, FIG. 4 is not necessarily to scale in allaspects. Like reference numerals in the Figures represent like features,and the description provided in relation to one Figure may apply to likefeatures in another Figure.

Referring to FIG. 4, the radius of the fan blade 23 (also referred to asthe radius of the tip 231 of the fan blade) is indicated by thedimension ‘Rfan’. The radius of the tip of the most downstreamcompressor blade 151 is indicated by the dimension ‘Rcomp’. Thecompression system radius ratio CSRR is thus defined as:

CSRR=Rfan/Rcomp

For the gas turbine engine 10, the value of CSRR may be in the rangesdefined herein, for example in the range of from 5 to 9, optionallyaround 5.2 to 8.5, optionally around 5.3 to 7.2, optionally around 5.3to 6.5.

The gas turbine engine 10 shown in FIG. 4 comprises a splitter 50 havinga leading edge at which the flow splits between the bypass flow B andthe core flow A. The radius of the leading edge of the splitter isindicated by the dimension ‘Rsplit’. The most downstream turbine blade191 has a radius indicated by the dimension ‘Rturb’. An engine coreradius ratio ECRR is defined as:

ECRR=Rturb/Rsplit

For the gas turbine engine 10, the compression system radius ratio(CSRR) divided by the engine core radius ratio (ECRR) may be in theranges defined herein, for example in the range of from 5.5 to 10,optionally 6 to 8. For the gas turbine engine 10, the compression systemblade ratio (defined elsewhere herein) divided by the engine core radiusratio may be in the ranges defined herein, for example in the range offrom 50 to 95, optionally 50 to 75. The ECRR itself may be, for example,in the range of from 0.75 to 1, for example 0.8 to 0.95.

The axial distance between the leading edge of the splitter 50 and theleading edge of the tip of the most downstream compressor blade 151 isindicated in FIG. 1 by the dimension ‘Xcomp’. A core compressor aspectratio CCAR is defined as:

CCAR=Xcomp/Rsplit

For the gas turbine engine 10, the compression system radius ratio(CSRR) divided by the core compressor aspect ratio (CCAR) may be in theranges defined herein, for example in the range of from 1.7 to 4.2,optionally 1.8 to 3.4. For the gas turbine engine 10, the compressionsystem blade ratio (defined elsewhere herein) divided by a corecompressor aspect ratio may be in the ranges defined herein, for examplein the range of from 15 to 50 The CCAR itself may be in the range offrom 2 to 3, for example 2.1 to 2.9, or 2.3 to 2.8.

A compression system speed ratio (CSSR) is defined as the ratio of therotational speed of the most downstream compressor blade 151 to therotational speed of the fan 23 at cruise conditions (the rotationalspeed of the most downstream compressor blade 151 being higher than therotational speed of the fan 23, of course). For the gas turbine engine10, the product of the compression system radius ratio and thecompression system speed ratio may be in the range of from 25 to 80, forexample in the range of from 35 to 65. For the gas turbine engine 10,the product of the compression system blade ratio and the compressionsystem speed ratio may be in the range of from 300 to 800, optionally320 to 750, optionally 325 to 700. The CSSR itself may be in the rangeof from 6.0 to 9.5, for example 6.5 to 9.0.

The fan blade has a height hfan. As indicated in FIG. 4 this is definedas the radius of the leading edge 232 at the tip 231 of the blade 23minus the radius of the point where the leading edge 232 intersects theradially inner gas-washed hub. Similarly, the blade height hcomp of themost downstream compressor blade 151 is defined as the radius of theleading edge at the tip of the blade minus the radius of the point wherethe leading edge intersects the radially inner gas-washed surface. Acompression system blade ratio CSBR is defined as:

CSBR=hfan/hcomp

For the gas turbine engine 10, the compression system blade ratio CSBRmay be in the ranges defined herein, for example in the range of from 45to 95, 50 to 75 or 55 to 70.

A core compressor pressure ratio (CCPR) is defined as the pressure (i.e.the mean total pressure) immediately downstream of the final rotor blade151 in the compressor (for example at the plane perpendicular to theaxial direction at the axial position indicated schematically byreference numeral 155 in FIG. 4) divided by the pressure (i.e. the meantotal pressure) immediately upstream of the first rotor blade 141 in thecore compressor (for example at the plane perpendicular to the axialdirection at the axial position indicated schematically by referencenumeral 145 in FIG. 4) at cruise conditions. In some arrangements, thecore compressor pressure ratio (which is defined at cruise conditions)may be in the range of from 34 to 60, for example 35, 36, 38 or 40 to55, for example 41 to 52 at cruise conditions.

A ratio of the core compressor aspect ratio divided by the corecompressor pressure ratio (i.e. CCAR/CCPR) may be in the ranges definedherein, for example in a range of from 0.03 to 0.09, for example in therange having a lower bound of any of 0.04, 0.045 or 0.05, and an upperbound of any of 0.06, 0.07, 0.08 or 0.085.

No compressor rotor blades other than the most upstream row of rotorblades 141 of the low pressure compressor 14 and the most downstream rowof compressor blades 151 of the high pressure compressor 15 are shown inFIG. 4. However, it will be appreciated that this is merely to assist inthe explanations provided herein, and that the low pressure compressor14 and the high pressure compressor 15 each comprise more than one rotorstage, each of which may have an associated stator stage. The totalnumber of rotor stages in the low pressure compressor 14 and the highpressure compressor 15 combined may be, for example, twelve, thirteen,or fourteen.

In a first arrangement of gas turbine engine 10, any one or more of thefollowing may apply:

-   -   the radius of the fan blade Rfan is 160 cm to 190 cm, the radius        of the tip of the most downstream compressor blade 151 is 27 cm        to 31 cm, and the CSRR in the range of from 5.3 to 7.7; by way        of non-limitative example, the radius of the fan blade Rfan is        175 cm and the radius of the tip of the most downstream        compressor blade 151 is 29 cm, giving a CSRR of 6.0    -   the radius of the most downstream turbine blade 191 is 65 cm to        80 cm, the radius of the leading edge of the splitter 50 is 70        cm to 90 cm, and the ECRR in the range of from 0.8 to 1; by way        of non-limitative example, the radius of the most downstream        turbine blade 191 is 75 cm and the radius of the leading edge of        the splitter 50 is 80 cm, giving an ECRR of 0.93    -   the axial distance between the leading edge of the splitter 50        and the leading edge of the tip of the most downstream        compressor blade 151 Xcomp is 180 cm to 225 cm, and the CCAR is        in the range of from 1.7 to 3.4; by way of non-limitative        example, the axial distance between the leading edge of the        splitter 50 and the leading edge of the tip of the most        downstream compressor blade 151 Xcomp is 195 cm, giving a CCAR        of 2.4    -   the fan blade height is 115 cm to 150 cm, the height of the most        downstream compressor blade is 1.9 cm to 2.3 cm, and the CSBR is        50 to 90; by way of non-limitative example, the fan blade height        is 125 cm, and the height of the most downstream compressor        blade is 2.1 cm, giving a CSBR of 60    -   at cruise conditions, the rotational speed of the fan 23 is 1300        rpm to 1800 rpm and the rotational speed of the most downstream        compressor blade 151 is 11000 rpm to 12000 rpm, and the CSSR is        in the range of from 6.5 to 9; by way of non-limitative example,        at cruise conditions, the rotational speed of the fan 23 is 1650        rpm and the rotational speed of the most downstream compressor        blade 151 is 12000 rpm, giving a CSSR of 7.3    -   at cruise conditions, the fan pressure ratio is 1.30 to 1.45,        the fan root pressure ratio is 1.18 to 1.30, the fan tip        pressure ratio is 1.30 to 1.45, and the core compressor pressure        ratio is 35 to 55; by way of non-limitative example, at cruise        conditions, the fan pressure ratio is 1.4, the fan root pressure        ratio is 1.25, the fan tip pressure ratio is 1.42, and the core        compressor pressure ratio is 44    -   at cruise conditions, the CCPR is 40 to 60, the ratio CCAR/CCPR        is 0.03 to 0.08, and the number of compressor rotor stages is 12        to 14; by way of non-limitative example, at cruise conditions,        the CCPR is 44, the ratio CCAR/CCPR is 0.055, and the number of        compressor rotor stages is 12.

Purely by way of example, the non-limitative examples referred to ineach of the bullets points above relating to a first arrangement mayrelate to the same engine.

In a second arrangement, any one or more of the following may apply:

-   -   the radius of the fan blade Rfan is 120 cm to 140 cm, the radius        of the tip of the most downstream compressor blade 151 is 20 cm        to 25 cm, and the CSRR is in the range of from 5.2 to 6.6; by        way of non-limitative example, the radius of the fan blade Rfan        is 130 cm and the radius of the tip of the most downstream        compressor blade 151 is 23 cm, giving a CSRR of 5.7    -   the radius of the most downstream turbine blade 191 Rturb is 40        cm to 60 cm, the radius of the leading edge of the splitter 50        Rsplit is 50 cm to 70 cm, and the ECRR is in the range of from        0.75 to 1.0; by way of non-limitative example, the radius of the        most downstream turbine blade 191 Rturb is 45 cm and the radius        of the leading edge of the splitter 50 Rsplit is 56 cm, giving        an ECRR of 0.80    -   the axial distance between the leading edge of the splitter 50        and the leading edge of the tip of the most downstream        compressor blade 151 Xcomp is 150 cm to 190 cm, and the CCAR is        in the range of from 2.2 to 3.8; by way of non-limitative        example, the axial distance between the leading edge of the        splitter 50 and the leading edge of the tip of the most        downstream compressor blade 151 Xcomp is 159 cm, giving a CCAR        of 2.8    -   the fan blade height is 75 cm to 100 cm, the height of the most        downstream compressor blade is 1.5 cm to 2.0 cm, and the CSBR is        45 to 75; by way of non-limitative example, the fan blade height        is 85 cm, and the height of the most downstream compressor blade        is 1.7 cm, giving a CSBR of 50    -   at cruise conditions, the rotational speed of the fan 23 is 2200        rpm to 2700 rpm, the rotational speed of the most downstream        compressor blade 151 is 14000 rpm to 17000 rpm, and the CSSR is        in the range of from 6 to 8; by way of non-limitative example,        at cruise conditions, the rotational speed of the fan 23 is 2500        rpm and the rotational speed of the most downstream compressor        blade 151 is 16000 rpm, giving a CSSR of 6.4    -   at cruise conditions, the fan pressure ratio is 1.30 to 1.45,        the fan root pressure ratio is 1.18 to 1.30, the fan tip        pressure ratio is 1.30 to 1.45, and the core compressor pressure        ratio is 35 to 55; by way of non-limitative example, at cruise        conditions, the fan pressure ratio is 1.4, the fan root pressure        ratio is 1.25, the fan tip pressure ratio is 1.42, and the core        compressor pressure ratio is 35    -   at cruise conditions, the CCPR is 34 to 50, the ratio CCAR/CCPR        is 0.05 to 0.09, and the number of compressor rotor stages is 12        to 14; by way of non-limitative example, at cruise conditions,        the CCPR is 35, the ratio CCAR/CCPR is 0.08, and the number of        compressor rotor stages is 12.

Purely by way of example, the non-limitative examples referred to ineach of the bullets points above relating to a second arrangement mayrelate to the same engine.

A further example of a feature that may be better optimized for gasturbine engines 10 according to the present disclosure compared withconventional gas turbine engines is the intake region, for example theratio between the intake length L and the fan diameter D. Referring toFIG. 1, the intake length L is defined as the axial distance between theleading edge of the intake and the leading edge 232 of the tip 231 ofthe fan blades, and the diameter D of the fan 23 is defined at theleading edge of the fan 23 (i.e. D=2×Rfan). Gas turbine engines 10according to the present disclosure, such as that shown by way ofexample in FIG. 1, may have values of the ratio L/D as defined herein,for example less than or equal to 0.45, for example 0.2 to 0.45. Thismay lead to further advantages, such as installation and/or aerodynamicbenefits.

The gas turbine engine 10 shown by way of example in the Figures maycomprise any one or more of the features described and/or claimedherein. For example, where compatible, such a gas turbine engine 10 mayhave any one or more of the features or values described herein of:CSRR; CCAR; ECRR; CSSR; CSBR; core compressor pressure ratio; CCAR/(corecompressor pressure ratio); CSRR/ECRR; CSRR/CCAR; CSRR*CSSR; CSBR/ECRR;CSBR/CCAR; CSBR*CSSR; number of compressor rotor stages; specificthrust; maximum thrust, turbine entry temperature; overall pressureratio; bypass ratio; fan diameter; fan rotational speed; fan hub to tipratio; fan pressure ratio; fan root pressure ratio; ratio between thefan root pressure ratio to the fan tip pressure ratio; fan tip loading;number of fan blades; construction of fan blades; and/or gear ratio.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

1. A gas turbine engine for an aircraft comprising: an engine corecomprising a turbine comprising a plurality of turbine blades, acompressor comprising a plurality of compressor blades, and a combustor;a fan comprising a plurality of fan blades; and a gearbox that receivesan input from a core shaft that is connected to at least a part of theturbine, the gearbox configured to drive the fan at a lower rotationalspeed than a rotational speed of the core shaft, wherein: a bypass ductis defined radially outside the core, with a leading edge of a splitterdefining a point at which flow splits into a core flow and a bypass flowduring operation of the gas turbine engine; a core compressor aspectratio is defined as a ratio of an axial distance between the leadingedge of the splitter and a leading edge of a tip of a downstream-mostcompressor blade to a radius of the leading edge of the splitter, theradius of the leading edge of the splitter being measured from acenterline axis of the gas turbine engine, the core compressor aspectratio being in a range of from 1.7 to 3.4; and an engine core radiusratio is defined as the ratio of the radius of the tip of the mostdownstream turbine blade in the engine to the radius of the leading edgeof the splitter, the engine core radius ratio being in a range of from0.8 and 0.98.
 2. The gas turbine engine according to claim 1, whereinthe core compressor aspect ratio is in a range of from 1.7 to 3.2. 3.The gas turbine engine according to claim 1, wherein the core compressoraspect ratio is in a range of from 1.8 to 3.0.
 4. The gas turbine engineaccording to claim 1, wherein the core compressor aspect ratio is in arange of from 1.9 to 2.9.
 5. The gas turbine engine according to claim1, wherein the core compressor aspect ratio is in a range of from 2.3 to2.8.
 6. The gas turbine engine according to claim 1, wherein the enginecore radius ratio is in a range of from 0.8 to 0.95.
 7. The gas turbineengine according to claim 1, wherein the engine core radius ratio is ina range of from 0.81 to 0.94.
 8. The gas turbine engine according toclaim 1, wherein the engine core radius ratio is in a range of from 0.82to 0.93.
 9. The gas turbine engine according to claim 1, wherein a corecompressor pressure ratio, defined as a pressure immediately downstreamof the downstream-most compressor blade divided by a pressureimmediately upstream of a forward-most compressor blade in thecompressor at cruise conditions, is in a range of from 33 to
 60. 10. Thegas turbine engine according to claim 9, wherein the ratio of the corecompressor aspect ratio divided by the core compressor pressure ratio isin a range of from 0.03 to 0.09.
 11. The gas turbine engine according toclaim 9, wherein the ratio of the core compressor aspect ratio dividedby the core compressor pressure ratio is in a range of from 0.04 to0.08.
 12. The gas turbine engine according to claim 1, wherein: a radiusof a tip of a fan blade of the plurality of fan blades is in a range offrom 120 cm to 210 cm, wherein the radius of the tip of the fan blade ismeasured from the centerline axis of the gas turbine engine; and/or thegearbox has a reduction ratio in a range of from 3.2 to 3.8, such thatthe rotational speed of the core shaft is in a range of from 3.2 to 3.8times the rotational speed of the fan.
 13. The gas turbine engineaccording to claim 9, wherein: the core compressor pressure ratio is ina range of from 36 to
 55. 14. The gas turbine engine according to claim1, wherein a compression system radius ratio defined as a ratio of aradius of a tip of a fan blade of the plurality of fan blades to aradius of the tip of the downstream-most compressor blade is in a rangeof from 5 to 7.5, the radius of the tip of the fan blade and the radiusof the tip of the downstream-most compressor blade each being measuredfrom the centerline axis of the gas turbine engine.
 15. The gas turbineengine according to claim 1, wherein a compression system speed ratio,defined as a ratio of a rotational speed of the downstream-mostcompressor blade to the rotational speed of the fan at cruiseconditions, is in a range of from 6 to
 9. 16. The gas turbine engineaccording to claim 1, wherein a fan pressure ratio, defined as a ratioof a mean total pressure of a flow at a fan exit to a mean totalpressure of a flow at a fan inlet at cruise conditions, is in a range offrom 1.35 to 1.43.
 17. The gas turbine engine according to claim 1,wherein: a fan root pressure ratio, defined as a ratio of a mean totalpressure of a flow at a fan exit that subsequently flows through theengine core to a mean total pressure of a flow at a fan inlet at cruiseconditions, is in a range of from 1.18 to 1.30.
 18. The gas turbineengine according to claim 1, wherein each fan blade of the plurality offan blades comprises a main body attached to a leading edge sheath, themain body and the leading edge sheath being formed using differentmaterials.
 19. The gas turbine engine according to claim 18, wherein theleading edge sheath material comprises titanium and/or the main bodymaterial comprises an aluminium alloy.
 20. The gas turbine engineaccording to claim 1, further comprising an intake that extends upstreamof the plurality of fan blades, wherein: an intake length L is definedas an axial distance between a leading edge of the intake and a leadingedge of a tip of a fan blade of the plurality of fan blades; a fandiameter D is a diameter of the fan at the leading edge of the tip ofthe fan blade; and a ratio L/D is in a range of from 0.2 to 0.45. 21.The gas turbine engine according to claim 1, wherein: the core shaft isa second core shaft; the compressor is a compressor section comprising afirst compressor and a second compressor; and the turbine is a turbinesection comprising a first turbine and a second turbine, wherein theengine core further comprises a first core shaft connecting the firstturbine to the first compressor; wherein the second core shaft connectsthe second turbine to the second compressor; the second turbine, thesecond compressor, and the second core shaft are arranged to rotate at ahigher rotational speed than a rotational speed of the first core shaft;and a number of compressor rotor stages is split between the firstcompressor and the second compressor.
 22. An aircraft comprising atleast one gas turbine engine according to claim 1.